This invention relates generally to gas turbine components, and more particularly to cooled turbine airfoils.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. In a turbofan engine, which typically includes a fan placed at the front of the core engine, a high pressure turbine powers the compressor of the core engine. A low pressure turbine is disposed downstream from the high pressure turbine for powering the fan. Each turbine stage commonly includes a stationary turbine nozzle followed in turn by a turbine rotor.
The turbine nozzle comprises a row of circumferentially side-by-side nozzle segments each including one or more stationary airfoil-shaped vanes mounted between inner and outer band segments for channeling the hot gas stream into the turbine rotor. Each of the vanes includes pressure and suction sidewalls that are connected at a leading edge and a trailing edge. The temperature distribution of a typical vane is such that the trailing edge is significantly hotter than the remainder of the airfoil. The temperature gradient created results in high compressive stress at the vane trailing edge, and the combination of high stresses and high temperatures generally results in the vane trailing edge being the life limiting location of the vane. Accordingly, in prior art vanes, the trailing edge portion is cooled using a source of relatively cool air, such as compressor discharge air, through a combination of internal convective cooling and film cooling. While this configuration increases the life of the vane, there remains a need for enhanced cooling of the trailing edge portion of turbine airfoils.